Geared turbofan engine with low pressure environmental control system for aircraft

ABSTRACT

An environmental control system includes a higher pressure tap at a higher pressure location in the main compressor section, and a lower pressure tap at a lower pressure location. The lower pressure location being at a lower pressure than the higher pressure location. The lower pressure tap communicates to a first passage leading to a downstream outlet and a compressor section of a turbocompressor. The higher pressure tap leads into a turbine section of the turbocompressor such that air in the higher pressure tap drives the turbine section to in turn drive the compressor section of the turbocompressor. A pylon includes a lowermost surface and the higher pressure tap does not extend above a plane including the lowermost surface. A combined outlet of the compressor section and the turbine section of the turbocompressor intermixes and passes downstream to be delivered to an aircraft use.

CROSS REFERENCE TO RELATED APPLICATION

This application claims priority to U.S. Provisional Application No.62/018,111 and U.S. Provisional Application No. 62/018,129 both filedJun. 27, 2014.

BACKGROUND OF THE INVENTION

This application relates to an environmental control system for anaircraft which taps both high and low pressure compressed air for useson an aircraft.

Environmental control systems are known, and associated with anaircraft. Typically, these systems tap air from a gas turbine engine onthe aircraft, and send it to the aircraft cabin, and other air uses onthe aircraft.

The systems typically tap low pressure compressed air from a lowerpressure compressor location, and higher pressure compressed air from ahigher pressure compressor location. The two are utilized at distincttimes during the operation of a gas turbine engine, dependent on theneeds, and the available air.

In the prior art, when the higher pressure air is tapped, it is at avery high temperature. Thus, cooling of the air must occur. It istypical that some form of intercooler or other heat exchanger isincluded.

In addition, the higher pressure compressed air has already beencompressed beyond the level of the lower pressure compressed air. Themore higher pressure compressed air that is diverted away from engineuses, the lower the efficiency of the engine.

SUMMARY

In a featured embodiment, a gas turbine engine assembly includes a fansection delivering air into a main compressor section. The maincompressor section compresses air and delivers air into a combustionsection. Products of combustion pass from the combustion section over aturbine section to drive the fan section and main compressor sections. Agearbox is driven by the turbine section to drive the fan section. Apylon supports the gas turbine engine. An environmental control systemincludes a higher pressure tap at a higher pressure location in the maincompressor section, and a lower pressure tap at a lower pressurelocation. The lower pressure location being at a lower pressure than thehigher pressure location. The lower pressure tap communicates to a firstpassage leading to a downstream outlet and a compressor section of aturbocompressor. The higher pressure tap leads into a turbine section ofthe turbocompressor such that air in the higher pressure tap drives theturbine section to in turn drive the compressor section of theturbocompressor. The pylon includes a lowermost surface and the higherpressure tap does not extend above a plane including the lowermostsurface. A combined outlet of the compressor section and the turbinesection of the turbocompressor intermixes and passes downstream to bedelivered to an aircraft use.

In another embodiment according to the previous embodiment, the gearboxprovides a gear reduction of at least about 2.0.

In another embodiment according to any of the previous embodiments, theturbine section includes a fan drive turbine that drives the gearbox andone of the main compressor sections.

In another embodiment according to any of the previous embodiments, themain compressor section includes a first compressor section and a secondcompressor section and the first compressor section includes at leastfour (4) stages and no more than seven (7) stages.

In another embodiment according to any of the previous embodiments,bleed air is taken from at least a fourth stage of the first compressorsection.

In another embodiment according to any of the previous embodiments, theturbine section includes a first turbine section driving a high pressurecompressor, an intermediate turbine section driving a low pressurecompressor and a third turbine section driving the fan section.

In another embodiment according to any of the previous embodiments, themain compressor section includes a first compressor section and a secondcompressor section and the first compressor section includes at leastthree (3) stages and no more than eight (8) stages.

In another embodiment according to any of the previous embodiments,bleed air is taken from at least a third stage of the first compressorsection.

In another embodiment according to any of the previous embodiments,includes a control valve in fluid communication with an inlet to thecompressor of the turbocompressor.

In another embodiment according to any of the previous embodiments,includes a valve disposed between low pressure tap and the compressorsection of the turbocompressor.

In another embodiment according to any of the previous embodiments,includes a sensor generating data indicative of a speed of the turbineof the turbocompressor.

In another embodiment according to any of the previous embodiments,includes a brake for controlling rotation of the turbine of theturbocompressor responsive to detection of an overspeed condition

In another featured embodiment, an environmental control system for anaircraft includes a higher pressure tap to be associated with a higherpressure location in a main compressor section associated with an engineof the aircraft, and a lower pressure tap to be associated with a lowerpressure location in the main compressor section. The lower pressurelocation being at a lower pressure than the higher pressure location.The lower pressure tap communicates to a first passage leading to adownstream outlet, and a compressor section of a turbocompressor. Thehigher pressure tap leads into a turbine section of the turbocompressorsuch that air in the higher pressure tap drives the turbine section toin turn drive the compressor section of the turbocompressor. The higherpressure tap is disposed below a plane including a lowermost surface ofa pylon supporting the main compressor section associated with theengine of the aircraft. A combined outlet of the compressor section andthe turbine section of the turbocompressor intermixes and passesdownstream to be delivered to an aircraft use.

In another embodiment according to the previous embodiment, a checkvalve is disposed within the first passage associated with the lowerpressure tap.

In another embodiment according to any of the previous embodiments, acontrol valve is positioned on the higher pressure tap, and may beclosed to drive air through the first passage associated with the lowerpressure tap, or to have air pass through the compressor section of theturbocompressor when the control valve is opened.

In another embodiment according to any of the previous embodiments, aredundant valve is provided to be closed if the control valve associatedwith the higher pressure tap fails.

In another embodiment according to any of the previous embodiments, theredundant valve is positioned to be downstream of a location at whichthe first passage and the combined outlet intermix into a commonconduit.

In another embodiment according to any of the previous embodiments,includes a control valve disposed between the low pressure tap and thecompressor section of the turbocompressor.

In another embodiment according to any of the previous embodiments,includes a sensor generating data indicative of a speed of the turbinesection of the turbocompressor.

In another embodiment according to any of the previous embodiments,includes a brake for controlling rotation of the turbine of theturbocompressor responsive to detection of an overspeed condition.

Although the different example have specific components shown in theillustrations, embodiments of this disclosure are not limited to thoseparticular combinations. It is possible to use some of the components orfeatures from one of the examples in combination with features orcomponents from another one of the examples.

These and other features disclosed herein can be best understood fromthe following specification and drawings, the following of which is abrief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1A schematically shows an embodiment of a gas turbine engine.

FIG. 1B schematically shows another gas turbine engine.

FIG. 2 shows an embodiment of an environmental control system for anaircraft.

FIG. 3 shows a schematic of the FIG. 2 system.

DETAILED DESCRIPTION

A gas turbine engine 210 is shown in FIG. 1A. As shown, the engine 210includes a fan 250 (which includes a plurality of fan blades 20), a maincompressor section 254 (which includes both a low pressure compressor256 and a high pressure compressor 258), a combustor 260, and a turbinesection 262 (which includes both a high pressure turbine 264 and a lowpressure turbine 266). The high pressure compressor 258 is driven, via afirst spool 268, by the high pressure turbine 264. The low pressurecompressor 256 is driven, via a second spool 270, by the low pressureturbine 266. Also driven by the low pressure turbine 266 are the fanblades 20 of the fan 250, which fan is coupled to the second spool 270via a geared architecture 272.

The fan section 250 drives air along a bypass flow path B while thecompressor section 254 draws air in along a core flow path C where airis compressed and communicated to a combustor section 260. In thecombustor section 260, air is mixed with fuel and ignited to generate ahigh pressure exhaust gas stream that expands through the turbinesection 262 where energy is extracted and utilized to drive the fansection 250 and the compressor section 254.

The second spool 270 generally includes an inner shaft 240 that connectsthe fan 250 and the low pressure (or first) compressor section 256 to alow pressure (or first) turbine section 266. The low pressure turbine266 is also referred to as the fan drive turbine as it drives the fan250 either directly or as is shown through the geared architecture 272.The inner shaft 240 drives the fan 250 through a speed change device,such as the geared architecture 272, to drive the fan 250 at a lowerspeed than the low speed spool 270. The high-speed spool 268 includes anouter shaft 242 that interconnects the high pressure (or second)compressor section 258 and the high pressure (or second) turbine section264. The inner shaft 240 and the outer shaft 242 are concentric androtate via the bearing systems disposed about the engine centrallongitudinal axis A.

Referring to FIG. 1B, another disclosed example gas turbine engine 215includes an intermediate or third spool 248. The engine 215 includesstructures similar to those disclosed and described with regard to theengine 210 shown in FIG. 1A such that like structures are provided withthe same reference numerals. The intermediate spool 248 includes anintermediate pressure turbine 246. The low pressure compressor 256 isdriven, via the intermediate spool 248 through an intermediate shaft 244coupled to the intermediate pressure turbine 246. The intermediate shaft244 is concentric with the inner shaft 240 of the second spool and theouter shaft 242 of the first spool 268. The low pressure turbine 266drives the fan blades 20 of the fan section 250. In this example, thelow pressure turbine 256 drives the inner shaft 240 to only drive thegeared architecture 272 that in turn drives the fan section 250. Itshould be appreciated, that the low pressure turbine 256 could alsodirectly drive the fan section without the speed reduction provided bythe geared architecture 272.

The disclosed gas turbine engines 210, 215 in one example arehigh-bypass geared aircraft engines. In a further example, the gasturbine engines 210, 215 each include a bypass ratio greater than aboutsix (6), with an example embodiment being greater than about ten (10).The example geared architecture 272 is an epicyclical gear train, suchas a planetary gear system, star gear system or other known gear system,with a gear reduction ratio of greater than about 2.0.

In the disclosed embodiments, the gas turbine engines 210, 215 include abypass ratio greater than about ten (10:1) and the fan diameter issignificantly larger than an outer diameter of the low pressurecompressor 256. It should be understood, however, that the aboveparameters are only exemplary of embodiments of a gas turbine engineincluding a geared architecture and that the present disclosure isapplicable to other gas turbine engines.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 250 is designed for a particularflight condition—typically cruise at about 0.8 Mach and about 35,000feet. The flight condition of 0.8 Mach and 35,000 ft., with the engineat its best fuel consumption—also known as “bucket cruise ThrustSpecific Fuel Consumption (‘TSFC’)”—is the industry standard parameterof pound-mass (lbm) of fuel per hour being burned divided by pound-force(lbf) of thrust the engine produces at that minimum point.

“Low fan pressure ratio” is the pressure ratio across the fan bladealone, without a Fan Exit Guide Vane (“FEGV”) system. The low fanpressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.50. In another non-limiting embodimentthe low fan pressure ratio is less than about 1.45.

“Low corrected fan tip speed” is the actual fan tip speed in ft/secdivided by an industry standard temperature correction of [(Tram °R)/(518.7° R)]^(0.5). The “Low corrected fan tip speed”, as disclosedherein according to one non-limiting embodiment, is less than about 1150ft/second.

The example low pressure compressor section 256 includes at least 4stages. In one disclosed embodiment the low pressure compressor 256includes seven (7) stages. In another disclosed embodiment the lowpressure compressor includes at least four (4) and up to seven (7)stages. In another disclosed embodiment, the example low pressurecompressor section 256 includes at least four (4) stages and up to abouteight (8) stages. In yet another disclosed embodiment the low pressurecompressor 256 includes eight (8) stages.

An environmental control system 30 for use on an aircraft receives airfrom portions of the compressor 254. In this example, the ECS system 30receives air from a portion of the low pressure compressor 256 and thehigh pressure compressor 258.

Referring to FIG. 2 with continued reference to FIGS. 1A and 1B, the ECS30 for use on an aircraft is illustrated. A high pressure compressionlocation 134 has a tap 34 as shown in FIG. 2. Another tap 32 is at alower pressure location 132. Locations 132 and 134 may both be withinthe high pressure compressor 258 or one may be in the lower pressurecompressor section 256. However, the tap 34 is downstream of the tap 32,and at a higher pressure location.

The compressor section 254, combustor 260 and the turbine section 262are disposed within a core cowling schematically indicated at 212. Thecore cowling 212 is disposed about the core engine features. The engines210, 215 are supported on an aircraft by a pylon 214 (shown in FIG. 2)that defines a lowermost surface 216 also referred to as a plane. Thelower plane 216 is the lowest extent of the pylon 214 toward the engine210, 215. The example ECS 30 including turbocompressor 42 (describedbelow) is disposed within the engine cavity defined within the corecowling 212.

Referring to FIG. 3 with continued reference to FIG. 2, the tap 32 leadsto first passage 36 having a check valve 38, and also into thecompressor section 54 of a turbocompressor 42. The high pressure tap 34leads into a turbine section 52 of the turbocompressor 42. The exits ofboth compressor section 54 and turbine section 52 of turbocompressor 42pass into a common outlet 44.

The outlet 44 merges with the first passage 36 and both pass through avalve 50 within a common outlet 37 leading to an aircraft use 152.

As shown in FIG. 3, the tap 32 alternatively leads to compressor section54 or into the first passage 36 leading to the combined outlet 37. Checkvalve 38 allows flow from tap 32 to the first passage 36 in a singledirection. It also provides some resistance to flow in that direction.The tap 34 leads through a modulating and shutoff valve 40 which can beopened or closed by a controller 41, shown schematically. Air from thehigher compressed location at tap 34 is expanded across the turbinesection 52 into the outlet 44. In one example higher pressure air isprovided from the high pressure compressor 258 to the tap 34.

The tap 34 communicates high pressure and temperature air to theturbocompressor 42 and does not pierce the plane 216 defined by thelowest portion of the pylon 214.

The high pressure and temperature air from the tap 34 drives the turbinesection 52 that drives the compressor section 54 to compress the airfrom the tap 32, and increase pressure of airflow into the combinedoutlet 37. Outlets from each of the turbine section 52 and thecompressor section 54 mix in the outlet 44, and pass to the combinedoutlet 37. When the compressor section 54 is being driven by the turbinesection 52, there is suction applied to the first passage 36 and the tap32, and thus check valve 38 will remain closed.

In one example, bleed air is taken from a fourth stage of the lowpressure compressor 256 and fed to the tap 32. The air from the tap 32is used generally exclusively under certain conditions when the heat tobe rejected is at a maximum. As an example, airflow will tend to passfrom tap 32 through the check valve 38 to the first passage 36 duringclimb and cruise. At such times the valve 40 is maintained closed tolimit the diversion of compressed air.

However, under certain conditions, as an example a descent, the valve 40is opened and the turbine section 52 is driven and air from tap 32passes to the compressor section 54. Expansion of higher temperature andpressure air from tap 34 through the turbine section 52 lowers itstemperature. Further, mixing it with the lower pressure compressed airfrom the tap 32, even when compressed to a higher pressure by compressorsection 54, may eliminate the need for a separate heat exchanger on theoutlet 44. The intermixed air may be at a useful temperature when itreaches the combined outlet 37. The amount of air from the two taps canbe varied to achieve this.

The valve 50 is a control valve which may be closed if the valve 40fails. At such times, it may be more desirable to supply no air to thesystem 152, then to have an open diversion from the tap 34.

A valve 100 is provided prior to the compressor section 54 and iscontrolled by the controller 41. The valve 100 is actuated to close offflow from the low pressure compressor 256 to control and modulate lowpressure airflow into the compressor 54.

A sensor 102 is provided that generates data indicative of turbine speedthat is sent by way of communication line 101 to the controller 41. Thesensor 102 is configured to provide information indicative of a turbineoverspeed condition. The controller 41 will actuate, and/or close valves100, 40 and 50 in a desired combination to prevent damage to the system.In one example, the controller 41 will receive information from thesensor 102 indicative of the onset, or actual overspeed condition of theturbine 52. The controller 41 utilizes at least data from the sensor102, along with other data available of engine operation to recognize acurrent or potential turbine speed condition that warrants shutdown orother remedial actions. The controller 41 may close the valve 40 toprevent flow of high pressure air that drives the turbine 52. A brake104 may also be employed to shutdown the turbine 52 if an overspeedcondition or other undesirable operating condition is detected orindicated.

The elimination of a required heat exchanger, and the use of less airfrom the higher compression location, is particularly valuable whencombined with a system incorporating a gear drive for the turbo fan,such as shown at 272 in FIGS. 1A and 1B.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of this disclosure. For that reason, the followingclaims should be studied to determine the true scope and content of thisdisclosure.

The invention claimed is:
 1. A gas turbine engine assembly comprising: afan section delivering air into a main compressor section, said maincompressor section compressing air and delivering air into a combustionsection, products of combustion passing from said combustion sectionover a turbine section to drive said fan section and main compressorsections, wherein a gearbox is driven by said turbine section to drivesaid fan section; a pylon supporting the gas turbine engine; anenvironmental control system including a higher pressure tap at a higherpressure location in said main compressor section, and a lower pressuretap at a lower pressure location, said lower pressure location being ata lower pressure than said higher pressure location; said lower pressuretap communicating to a first passage leading to a downstream outlet anda compressor section of a turbocompressor; said higher pressure tapleading into a turbine section of said turbocompressor such that air insaid higher pressure tap drives said turbine section to in turn drivesaid compressor section of said turbocompressor, wherein the pylonincludes a lowermost surface and the higher pressure tap does not extendabove a plane including the lowermost surface; a combined outlet of saidcompressor section and said turbine section of said turbocompressorintermixing and passing downstream to be delivered to an aircraft use; acheck valve disposed within the first passage between said low pressuretap and said downstream outlet; and a control valve disposed between thelow pressure tap and the compressor section of the turbocompressor forcontrolling airflow from the low pressure tap into an inlet of thecompressor section of the turbocompressor.
 2. The gas turbine engineassembly as recited in claim 1, wherein the turbine section includes afan drive turbine that drives the gearbox and one of the main compressorsections.
 3. The gas turbine engine assembly as recited in claim 1,wherein the main compressor section includes a first compressor sectionand a second compressor section and the first compressor sectionincludes at least four (4) stages and no more than seven (7) stages. 4.The gas turbine assembly engine as recited in claim 3, wherein bleed airis taken from at least a fourth stage of the first compressor section.5. The gas turbine engine assembly as recited in claim 1, wherein theturbine section includes a first turbine section driving a high pressurecompressor, an intermediate turbine section driving a low pressurecompressor and a third turbine section driving the fan section.
 6. Thegas turbine engine assembly as recited in claim 1, wherein the maincompressor section includes a first compressor section and a secondcompressor section and the first compressor section includes at leastthree (3) stages and no more than eight (8) stages.
 7. The gas turbineengine assembly as recited in claim 6, wherein bleed air is taken fromat least a third stage of the first compressor section.
 8. The gasturbine engine assembly as recited in claim 1, including a sensor forgenerating data indicative of a speed of the turbine section of theturbocompressor.
 9. The gas turbine engine assembly as recited in claim8, including a brake for controlling rotation of the turbine section ofthe turbocompressor responsive to detection of an overspeed condition.10. An environmental control system for an aircraft comprising: a higherpressure tap to be associated with a higher pressure location in a maincompressor section associated with an engine of the aircraft, and alower pressure tap to be associated with a lower pressure location insaid main compressor section, said lower pressure location being at alower pressure than said higher pressure location; said lower pressuretap communicating to a first passage leading to a downstream outlet, anda compressor section of a turbocompressor; said higher pressure tapleading into a turbine section of said turbocompressor such that air insaid higher pressure tap drives said turbine section to in turn drivesaid compressor section of said turbocompressor, wherein said higherpressure tap is disposed below a plane including a lowermost surface ofa pylon supporting the main compressor section associated with theengine of the aircraft; a combined outlet of said compressor section andsaid turbine section of said turbocompressor intermixing and passingdownstream to be delivered to an aircraft use; a check valve is disposedwithin said first passage associated with said lower pressure tap; and acontrol valve disposed between the lower pressure tap and the compressorsection of the turbocompressor.
 11. The environmental control system asrecited in claim 10, including a sensor for generating data indicativeof a speed of the turbine section of the turbocompressor.
 12. Theenvironmental control system as recited in claim 11, including a brakefor controlling rotation of the turbine section of the turbocompressorresponsive to detection of an overspeed condition.
 13. An environmentalcontrol system for an aircraft comprising: a higher pressure tap to beassociated with a higher pressure location in a main compressor sectionassociated with an engine of the aircraft, and a lower pressure tap tobe associated with a lower pressure location in said main compressorsection, said lower pressure location being at a lower pressure thansaid higher pressure location; said lower pressure tap communicating toa first passage leading to a downstream outlet, and a compressor sectionof a turbocompressor; said higher pressure tap leading into a turbinesection of said turbocompressor such that air in said higher pressuretap drives said turbine section to in turn drive said compressor sectionof said turbocompressor, wherein said higher pressure tap is disposedbelow a plane including a lowermost surface of a pylon supporting themain compressor section associated with the engine of the aircraft; acombined outlet of said compressor section and said turbine section ofsaid turbocompressor intermixing and passing downstream to be deliveredto an aircraft use; and a control valve positioned on said higherpressure tap, and may be closed to drive air through said first passageassociated with said lower pressure tap, or to have air pass throughsaid compressor section of said turbocompressor when said control valveis opened.
 14. The environmental control system as recited in claim 13,wherein a redundant valve is provided to be closed if said control valveassociated with said higher pressure tap fails.
 15. The environmentalcontrol system as recited in claim 14, wherein said redundant valve ispositioned to be downstream of a location at which said first passageand said combined outlet intermix into a common conduit.